Crossover cooling flow for multi-engine systems

ABSTRACT

A multi-engine system includes a first gas turbine engine that includes a first compressor and a first turbine. The multi-engine system may further include a second gas turbine engine that has a second compressor and a second turbine. Still further, the multi-engine system may include a first crossover cooling network configured to route a first crossover airflow from the first compressor of the first gas turbine engine to the second turbine of the second gas turbine engine and a second crossover cooling network configured to route a second crossover airflow from the second compressor of the second gas turbine engine to the first turbine of the first gas turbine engine.

FIELD

The present disclosure relates to multi-engine systems, and morespecifically to inter-engine cooling flow.

BACKGROUND

Many systems utilize multiple engines to provide power. For example, arotorcraft (e.g., a helicopter) may utilize two or more engines to drivethe rotors and otherwise provide power to the rotorcraft. The multipleengines of these systems are often operated symmetrically. That is,these multi-engine systems are often controlled so that each engine ofthe multiple engines functions at substantially the same operatingparameters. However, overall fuel consumption may be reduced by, forexample, shutting down one of the engines while leaving one of theengines operating. Such asymmetric operation, however, has variouscontrol challenges.

SUMMARY

In various embodiments, the present disclosure provides a multi-enginesystem. The multi-engine system may include a first gas turbine enginethat includes a first compressor and a first turbine. The multi-enginesystem may further include a second gas turbine engine that has a secondcompressor and a second turbine. Still further, the multi-engine systemmay include a first crossover cooling network configured to route afirst crossover airflow from the first compressor of the first gasturbine engine to the second turbine of the second gas turbine engineand a second crossover cooling network configured to route a secondcrossover airflow from the second compressor of the second gas turbineengine to the first turbine of the first gas turbine engine.

In various embodiments, the first gas turbine engine has a first powerturbine, the second gas turbine engine has a second power turbine, themulti-engine system is a rotorcraft with a main rotor gearbox, and themain rotor gearbox is mechanically coupled to both the first powerturbine and the second power turbine. In various embodiments, the firstcrossover cooling network extends from the first compressor to a secondvane row of the second turbine and the second crossover cooling networkextends from the second compressor to a first vane row of the firstturbine. For example, the second vane row may be a second forward-mostvane row of the second turbine and the first vane row may be a firstforward-most vane row of the first turbine.

In various embodiments, the second forward-most vane row of the secondturbine may comprise a plurality of second vanes, wherein each secondvane of the plurality of second vanes defines a second leading edgechamber and a second body chamber aft of the second leading edgechamber. The first crossover cooling network may be configured to routethe first crossover airflow to the second body chamber. Further, thefirst forward-most vane row of the first turbine may comprise aplurality of first vanes, wherein each first vane of the plurality offirst vanes defines a first leading edge chamber and a first bodychamber aft of the first leading edge chamber. The second crossovercooling network may be configured to route the second crossover airflowto the first body chamber.

In various embodiments, the first gas turbine engine includes a firstintra-engine cooling network configured to route a first residentairflow from forward of the first turbine to the first leading edgechamber of the plurality of first vanes of the first forward-most vanerow of the first turbine. In various embodiments, the second gas turbineengine includes a second intra-engine cooling network configured toroute a second resident airflow from forward of the second turbine tothe second leading edge chamber of the plurality of second vanes of thesecond forward-most vane row of the second turbine.

In various embodiments, in response to the multi-engine system operatingin an intermediate rated power mode, the first crossover airflow isbetween 5% and 20% of a first total compressor flow through the firstcompressor and the second crossover airflow is between 5% and 20% of asecond total compressor flow through the second compressor. In variousembodiments, in response to the multi-engine system operating in anintermediate rated power mode, the first crossover airflow is 10% of thefirst total compressor flow through the first compressor and the secondcrossover airflow is 10% of the second total compressor flow through thesecond compressor. In various embodiments, in response to themulti-engine system operating in an intermediate rated power mode, thefirst total compressor flow is 100% of a first compressor inletcorrected flow capacity of the first compressor and the second totalcompressor flow is 100% of a second compressor inlet corrected flowcapacity of the second compressor.

In various embodiments, in response to the multi-engine system operatingin an asymmetric cruise mode, the first crossover airflow is between 5%and 20% of a first total compressor flow through the first compressorand the second crossover airflow is 0% of a second total compressor flowthrough the second compressor. In various embodiments, in response tothe multi-engine system operating in an asymmetric cruise mode, thefirst crossover airflow is 10% of the first total compressor flowthrough the first compressor and the second crossover airflow is 0% ofthe second total compressor flow through the second compressor. Invarious embodiments, in response to the multi-engine system operating inan asymmetric cruise mode, the first total compressor flow is 99% of afirst compressor inlet corrected flow capacity of the first compressorand the second total compressor flow is 40% of a second compressor inletcorrected flow capacity of the second compressor. In variousembodiments, in response to the multi-engine system operating in anasymmetric cruise mode, the second gas turbine engine is operating at afuel rate that is lower than self-sustaining idle parameters of thesecond gas turbine engine when operating independent of the firstcrossover airflow from the first gas turbine engine. In variousembodiments, an electric load, via a generator, is applied to the secondgas turbine engine.

In various embodiments, in response to the multi-engine system operatingin a one-engine-inoperable mode, the first crossover airflow is 10% of afirst total compressor flow through the first compressor and the secondcrossover airflow is 0% of a second total compressor flow through thesecond compressor. In various embodiments, in response to themulti-engine system operating in a one-engine-inoperable mode, the firsttotal compressor flow is 105% of a first compressor inlet corrected flowcapacity of the first compressor and the second total compressor flow is0% of a second compressor inlet corrected flow capacity of the secondcompressor.

In various embodiments, the first crossover cooling network includes atleast one of a first check valve configured to prevent backflow of thefirst crossover airflow and a first controlled valve configured tocontrol the first crossover airflow and the second crossover coolingnetwork comprises at least one of a second check valve configured toprevent backflow of the second crossover airflow and a second controlledvalve configured to control the second crossover airflow.

Also disclosed herein, according to various embodiments, is a method ofoperating a multi-engine system. The method may include, in response tothe multi-engine system operating in an asymmetric cruise mode, routinga portion of a total compressor flow of a compressor of a first gasturbine engine through a crossover cooling network that extends from thecompressor of the first gas turbine engine to a second gas turbineengine. The method may also include operating the second gas turbineengine at a fuel rate that is lower than a self-sustaining idle fuelrate of the second gas turbine engine if the second gas turbine enginewere operating independent of the portion of the total compressor flowof the compressor of the first gas turbine engine. In variousembodiments, during operation of the multi-engine system in theasymmetric cruise mode, overall power generation is the same as, butoverall fuel consumption is lower than, if the multi-engine system wereoperating under a symmetric cruise mode.

The forgoing features and elements may be combined in variouscombinations without exclusivity, unless expressly indicated hereinotherwise. These features and elements as well as the operation of thedisclosed embodiments will become more apparent in light of thefollowing description and accompanying drawings.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a schematic cross-sectional view of an exemplary gas turbineengine, in accordance with various embodiments;

FIG. 2 is a schematic view of a multi-engine system having crossovercooling flow networks, in accordance with various embodiments;

FIG. 3 is a graph showing operating lines of a multi-engine system, inaccordance with various embodiments;

FIG. 4A is a schematic view of a multi-engine system operating in anasymmetric cruise mode, in accordance with various embodiments;

FIG. 4B is a graph showing operation of a multi-engine system in anasymmetric cruise mode, in accordance with various embodiments;

FIG. 5A is a schematic view of a multi-engine system operating in anintermediate rated power mode, in accordance with various embodiments;

FIG. 5B is a graph showing operation of a multi-engine system in anintermediate rated power mode, in accordance with various embodiments;

FIG. 6A is a schematic view of a multi-engine system operating in aone-engine-inoperable mode, in accordance with various embodiments;

FIG. 6B is a graph showing operation of a multi-engine system in aone-engine-inoperable mode, in accordance with various embodiments; and

FIG. 7 is a schematic flow chart diagram of a method of operating amulti-engine system, in accordance with various embodiments.

The subject matter of the present disclosure is particularly pointed outand distinctly claimed in the concluding portion of the specification. Amore complete understanding of the present disclosure, however, may bestbe obtained by referring to the detailed description and claims whenconsidered in connection with the drawing figures, wherein like numeralsdenote like elements.

DETAILED DESCRIPTION

The detailed description of exemplary embodiments herein makes referenceto the accompanying drawings, which show exemplary embodiments by way ofillustration. While these exemplary embodiments are described insufficient detail to enable those skilled in the art to practice thedisclosure, it should be understood that other embodiments may berealized and that logical changes and adaptations in design andconstruction may be made in accordance with this disclosure and theteachings herein without departing from the spirit and scope of thedisclosure. Thus, the detailed description herein is presented forpurposes of illustration only and not of limitation.

As used herein, “aft” refers to the direction associated with theexhaust (e.g., the back end) of a gas turbine engine. As used herein,“forward” refers to the direction associated with the intake (e.g., thefront end) of a gas turbine engine. A first component that is “radiallyoutward” of a second component means that the first component ispositioned at a greater distance away from the engine centrallongitudinal axis than the second component. A first component that is“radially inward” of a second component means that the first componentis positioned closer to the engine central longitudinal axis than thesecond component. In the case of components that rotatecircumferentially about the engine central longitudinal axis, a firstcomponent that is radially inward of a second component rotates througha circumferentially shorter path than the second component. Theterminology “radially outward” and “radially inward” may also be usedrelative to references other than the engine central longitudinal axis.For example, a first component of a combustor that is radially inward orradially outward of a second component of a combustor is positionedrelative to the central longitudinal axis of the combustor. The term“axial,” as used herein, refers to a direction along or parallel to theengine central longitudinal axis.

Unless otherwise indicated, the terms “first,” “second,” etc. are usedherein merely as labels, and are not intended to impose ordinal,positional, or hierarchical requirements on the items to which theseterms refer. Moreover, reference to, e.g., a “second” item does notrequire or preclude the existence of, e.g., a “first” or lower-numbereditem, and/or, e.g., a “third” or higher-numbered item.

Disclosed herein, according to various embodiments, is a multi-enginesystem comprising inter-engine cooling flows. While numerous details andembodiments herein reference a rotorcraft (e.g., a helicopter), thescope of the disclosure is not necessarily limited to rotorcraft, andthus the multi-engine system may be implemented in power generationapplications, vehicles, other aircraft, etc.

Referring now to FIG. 1 and according to various embodiments, anexemplary gas turbine engine 10 is shown. As described in greater detailbelow, the multi-engine system 100 (FIG. 2) may include multipleidentical gas turbine engines, such as exemplary gas turbine engine 10,or the multi-engine system 100 may include engines that are notidentical. In various embodiments, the gas turbine engine 10 is aturboshaft engine, such as can be used in an aircraft application, suchas a helicopter. In various embodiments, the gas turbine engine 10comprises a gas generator 12 and/or a power turbine 14 arranged in aseries flow with an upstream inlet proximate the gas generator 12 and adownstream exhaust proximate the power turbine 14.

In various embodiments, the gas generator 12 includes a compressorsection 16 where air is compressed/pressurized, a combustor section 18downstream of the compressor section 16 where the compressed air ismixed with fuel and ignited to generate hot combustion gases, and aturbine section 20 downstream of the combustor section 18 for extractingpower from the hot combustion gases, such as by causing the blades of aturbine to rotate.

In various embodiments, the gas generator 12 further comprises amulti-spool coaxially nested configuration, including a low pressurespool 22 and a high pressure spool 24. In various embodiments, the lowpressure spool 22 and the high pressure spool 24 operate in differentdirections, as well as at different pressures, speeds, and/ortemperatures. In various embodiments, a low pressure turbine 26 ismounted to the low pressure spool 22 to drive a low pressure compressor28, and a high pressure turbine 30 is mounted to the high pressure spool24 to drive a high pressure compressor 32. As used herein, “lowpressure” components generally experience lower pressures thancorresponding “high pressure” components when the gas turbine engine 10operates.

In various embodiments, the power turbine 14 comprises a turbine 34(e.g., a “power turbine”) mounted to a turbine spool 36. In operation,the gas generator 12 generates combustion gases that impart torque tothe turbine spool 36 through the turbine 34. In various embodiments, theturbine spool 36 drives a load system 38, such as an electricalgenerator, power turbine, propeller, rotor, pump system, etc.

In various embodiments, the multi-engine system 100 (FIG. 2) and/or eachgas turbine engine 10 may have a controller 40 configured to control themulti-engine system 100 (FIG. 2) and/or each gas turbine engine 10.Computer-based system program instructions and/or processor instructionsmay be loaded onto a tangible, non-transitory computer readable mediumhaving instructions stored thereon that, in response to execution by aprocessor, cause the processor to perform various operations. The term“non-transitory” is to be understood to remove only propagatingtransitory signals per se from the claim scope and does not relinquishrights to all standard computer-readable media that are not onlypropagating transitory signals per se. Stated another way, the meaningof the term “non-transitory computer-readable medium” and“non-transitory computer-readable storage medium” should be construed toexclude only those types of transitory computer-readable media that werefound in In re Nuijten to fall outside the scope of patentable subjectmatter under 35 U.S.C. § 101.

In various embodiments, the controller 40 comprises a Full-AuthorityDigital Engine Control (FADEC) system for use with an aircraft gasturbine engine 10. In various embodiments, the controller 40 includesone or more processors 42 and one or more tangible, non-transitorymemories 44 configured to implement digital or programmatic logic. Invarious embodiments, for example, the one or more processors 42 compriseone or more of an application specific integrated circuit (ASIC),digital signal processor (DSP), field programmable gate array (FPGA),general purpose processor, and/or other programmable logic device,discrete gate, transistor logic, or discrete hardware components, or anyvarious combinations thereof and/or the like, and the one or moretangible, non-transitory memories 44 store instructions that areimplemented by the one or more processors 42 for performing variousfunctions, such as the systems and methods of the inventive arrangementsdescribed herein.

In various embodiments, and with reference to FIGS. 2 and 3, themulti-engine system 100 is provided. The multi-engine system 100generally includes a first gas turbine engine 110 and a second gasturbine engine 120. Each gas turbine engine 110, 120, which may be thesame as or similar to gas turbine engine 10 described above withreference to FIG. 1, has a compressor 111, a diffuser-combustor 112, aturbine 113, etc., according to various embodiments. That is, the firstgas turbine engine 110 may have a first compressor 111, a firstdiffuser-combustor 112, and a first turbine 113 and the second gasturbine engine 120 may have a second compressor 121, a seconddiffuser-combustor 122, and a second turbine 123. The first and secondgas turbine engines 110, 120 may each include a power turbine—first gasturbine engine 110 may include a first power turbine 118 and the secondgas turbine engine may include a second power turbine 128—and thesepower turbines may both be mechanically coupled, for example, to a mainrotor gearbox of a rotorcraft.

The multi-engine system 100 may further include a means for routingcooling flow between the gas turbine engines. For example, themulti-engine system 100 may include a first crossover cooling network131 and a second crossover cooling network 132. Generally, thesecrossover cooling networks 131, 132 are configured to createinter-engine cooling flow between the first and second gas turbineengines 110, 120. That is, the first crossover cooling network 131 maybe configured to route a first crossover airflow from the firstcompressor 111 of the first gas turbine engine 110 to the second turbine123 of the second gas turbine engine 120 and the second crossovercooling network 132 may be configured to route a second crossoverairflow from the second compressor 121 of the second gas turbine engine120 to the first turbine 113 of the first gas turbine engine 110. Asdescribed in greater detail below with reference to the variousoperating modes of the multi-engine system 100, the dual crossovercooling networks 131, 132 enable the inter-engine transfer of flowenergy (e.g., enthalpy) in order to enable asymmetric operation of thegas turbine engines 110, 120, which may result in overall improved fuelefficiency of the multi-engine system 100. While numerous details andembodiments are included herein pertaining to a configuration of two gasturbine engines 110, 120, the multi-engine system 100 may include morethan two engines (e.g., three or more).

In various embodiments, the first and second crossover cooling networks131, 132 each comprises a series of tubes, pipes, channels, chambers,plenums, etc., for directing cooling airflow from the compressor of onegas turbine engine to the turbine of the other gas turbine engine. Thefirst and second crossover airflows that flow through the first andsecond crossover cooling networks 131, 132, respectively, may be apercentage of total compressor flow. That is, the first crossoverairflow being routed via the first crossover cooling network 131 maycomprise between about 5% and about 20% of a first total compressor flowthrough the first compressor 111 and the second crossover airflow beingrouted via the second crossover cooling network 132 may similarlycomprises between about 5% and about 20% of a second total compressorflow through the second compressor 121. In various embodiments, thefirst crossover airflow is about 10% of the first total compressor flowthrough the first compressor 111 and the second crossover airflow isabout 10% of the second total compressor flow through the secondcompressor 121. As used in this context, the term “about” refers to plusor minus 2%.

As used herein, the term “crossover airflow” refers to the cooling airfrom the compressors of one of the gas turbine engines that isconfigured to crossover (i.e., flow) from one engine to the other.Accordingly, the term “first crossover airflow” refers to the air beingdirected inter-engine (e.g., from the first gas turbine engine 110 tothe second gas turbine engine 120) via the first crossover coolingnetwork 131 and the term “second crossover airflow” refers to the airbeing directed inter-engine (e.g., from the second gas turbine engine120 to the first gas turbine engine 110) via the second crossovercooling network 132. Thus, while each of the first and second crossovercooling networks 131, 132 shown in FIG. 2 has a split that allows thecooling air to flow to the other engine or to flow directly to its ownengine's turbine, the term “crossover airflow” refers only to theportion that that actually transfers between engines and does not referto the flow of air that returns back to the same engine from which itwas diverted. Accordingly, each crossover cooling network 131, 132 mayinitially (e.g., at or near the inlet of the crossover cooling networkwhere cooling air is diverted from the compressor) have between about10% and about 40% of the total compressor flow being directedthere-through, which may be divided downstream so that between about 5%and about 20% (relative to the total compressor flow) flows to the othergas turbine engine (the “crossover airflow), with the other portion(5%-20%) flowing to the turbine of the same engine from which it wasdiverted. In various embodiments, each crossover cooling network 131,132 may initially (e.g., at or near the inlet of the crossover coolingnetwork where cooling air is diverted from the compressor) have betweenabout 15% and about 30% of the total compressor flow being directedthere-through. In various embodiments, each crossover cooling network131, 132 may initially (e.g., at or near the inlet of the crossovercooling network where cooling air is diverted from the compressor) haveabout 10% of the total compressor flow being directed there-through. Asused in this context only, the term “about” refers to plus or minus 2%.Accordingly, in various embodiments, vanes of the turbine of one of thegas turbine engines may be configured to be supplied with coolingairflow from its own compressor and from the other gas turbine engine.Additional details pertaining to the relative flows of cooling airfloware included below with reference to exemplary operating modes of themulti-engine system 100.

In various embodiments, the first crossover cooling network 131 extendsfrom the first compressor 111 of the first gas turbine engine 110 to asecond vane row 124 (no ordinal or positional meaning is intended forthe term “second” in this context) of the second turbine 123 of thesecond gas turbine engine 120. Similarly, the second crossover coolingnetwork 132 may extend from the second compressor 121 of the second gasturbine engine 120 to a first vane row 114 of the first turbine 113 ofthe first gas turbine engine 110. In various embodiments, the secondvane row 124 is a second forward-most vane row of the second turbine 123and the first vane row 114 is a first forward-most vane row of the firstturbine 113. Said differently, the vanes to which the crossover airflowsare directed may be the vanes immediately downstream (e.g., aft) of thediffuser-combustor 112, 122 which are subjected to comparatively highertemperatures, according to various embodiments.

In various embodiments, the first forward-most vane row (e.g., 114) ofthe first turbine 113 comprises a plurality of first vanes, wherein eachfirst vane of the plurality of first vanes defines a first leading edgechamber 115 and a first body chamber 116 that is aft of the firstleading edge chamber 115. The second crossover cooling network 132 maybe configured to route the second crossover airflow to the first bodychamber 116 and may not direct the second crossover airflow to the firstleading edge chamber 115. In such a configuration, the leading firstedge chamber 115 may be supplied with a first resident airflow via afirst intra-engine cooling network 133. That is, the first leading edgechamber 115, which may have an array of cooling holes (e.g., alsoreferred to as a “showerhead”-type array of cooling holes) along oradjacent the leading edge of the airfoil shape of the first vanes, maybe fluidly isolated from the second crossover flow and may instead besupplied by its own engine's cooling flow, referred to herein as firstresident airflow, via the first intra-engine cooling network 133.Correspondingly, the second forward-most vane row (e.g., 124) of thesecond turbine 123 comprises a plurality of second vanes, wherein eachsecond vane of the plurality of second vanes defines a second leadingedge chamber 125 and a second body chamber 126 that is aft of the secondleading edge chamber 125. The first crossover cooling network 131 may beconfigured to route the first crossover airflow to the second bodychamber 126 and may not direct the first crossover airflow to the secondleading edge chamber 125. In such a configuration, the second leadingedge chamber 125 may be supplied with a second resident airflow via asecond intra-engine cooling network 134. That is, the second leadingedge chamber 125, which may have “showerhead-type” cooling holes alongor adjacent the leading edge of the airfoil shape of the second vanes,may be fluidly isolated from the first crossover flow and may instead besupplied by its own engine's cooling flow, referred to herein as secondresident airflow, via the second intra-engine cooling network 134. Eachof the gas turbine engines 110, 120 may also have other intra-enginecooling networks, and/or the described cooling networks may be routedthrough various portions of the respective engines.

In various embodiments, and with reference to FIG. 3, a conventionaloperating line 350 of a gas turbine engine is provided, as well as amodified operating line 355 of a gas turbine engine that has beenconfigured to provide crossover flow to another gas turbine engine. Invarious embodiments, the gas turbine engines 110, 120 of themulti-engine system 100 have been designed and/or configured to haveextra cooling flow capacity (as indicated by the modified operating line355). The conventional operating line 350 shows a conventional cruiseoperating point as well as a conventional self-sustaining idle pointthat represents the lowest operating conditions of a gas turbine enginewhile still maintaining a self-sufficient idle.

In various embodiments, and with reference to FIGS. 4A and 4B, anasymmetric operating mode of the multi-engine system 100 is provided.The asymmetric operating mode of FIGS. 4A and 4B may also be referred toas an asymmetric cruise mode. In the asymmetric cruise mode, the firstgas turbine engine 110 may be generally configured to increase itsoperation and the second gas turbine engine 120 may be generallyconfigured to decrease its operation. Said differently, the first gasturbine engine 110 may be configured to operate at a higher power ratingwhile the second gas turbine engine 120 may be configured to operate ata lower power rating. In various embodiments, this lower power rating ofthe second gas turbine engine 120 is lower than the aforementionedself-sustaining idle point of a conventional gas turbine engine (e.g.,lower than the self-sustaining idle point of the second gas turbineengine with no crossover flow). That is, as schematically depicted andrepresented by arrows 419 and 429 in FIG. 4A, and as it corresponds tothe x-axis of the graph of FIG. 4B, the first gas turbine engine 110 maybe configured to operate with a first total compressor flow that isoperating at or near full capacity. That is, the first gas turbineengine 110 may operate with a first total compressor flow that isbetween about 90% and 100% (e.g., 99%) of a first compressor inletcorrected flow capacity of the first compressor 111 while the second gasturbine engine 120 may be configured to operate with a second totalcompressor flow that is less than or equal to 40% of a second compressorinlet corrected flow capacity of the second compressor 121.

As seen in the graph in FIG. 4B, the operating point of the second gasturbine engine 120 in the asymmetric cruise mode may be below what wouldotherwise be the minimal operating power of the second gas turbineengine 120 to maintain self-sustaining idle. This lower operating pointis possible because of the first crossover airflow from the first gasturbine engine 110 to the second turbine 123 of the second gas turbineengine 120. This first crossover airflow provides flow energy (e.g.,enthalpy) to the second gas turbine engine 120, thereby enabling thefuel rate of the second gas turbine engine 120 to be lower than wouldotherwise be possible because the first crossover airflow augments andsupplements the operation of the second gas turbine engine 120. Invarious embodiments, the first gas turbine engine 110 operates moreefficiently at the elevated/higher operating point, and thus the overallfuel consumption rate of the multi-engine system 100 operating in theasymmetric cruise mode described herein may be lower than if both gasturbine engines 110, 120 were operating under a symmetric/conventionalcruise mode. In various embodiments, the second gas turbine engine 120,even at this lower operating point, may have an electric load appliedthereto, such as via a generator mechanically coupled to one or more ofthe turbine sections of the gas turbine engine 120.

In various embodiments, the first crossover airflow flowing through thefirst crossover cooling network 131 is between about 5% and about 20%,as mentioned above, of the first total compressor flow through the firstcompressor 111 while there is no flow through the second crossovercooling network 132 (e.g., the second crossover airflow is 0). Theone-directional crossover flow from the first gas turbine engine 110 tothe second gas turbine engine 120 may be passively achieved due to thedifference of the pressures between the cooling airflow from therespective compressors 111, 121 of the gas turbine engines 110, 120. Invarious embodiments, the crossover cooling networks 131, 132 may includeone or more valves, such as a check valve to prevent backflow of thecrossover airflows and/or a controlled valve to control the flow of thecrossover airflows. Accordingly, the one directional crossover flow(from the engine at the higher operating power to the engine at thelower operating power) may be actively controlled via a controlled valve(e.g., one or both of the valves 135 shown in FIG. 4A).

Below is a table that shows one example of the contrasting operatingvariables of a multi-engine system operating in a conventional cruisemode versus the asymmetric cruise mode described herein and enabled bythe crossover airflow:

Cruise For One hour Conventional Axisymmetric Hybrid 100% Power = EngineEngine Engine Engine 7000 shp No. 1 No. 2 No. 1 No. 2 Turbine 3100 31003260 2050 Temperature Thermodynamic 39.3% 39.3% 40.5% 33.2% EfficiencySFC 0.353 0.353 0.335 0.410 % Flow 60 60 94 18 % Cross Bleed 0 0 — 9.4Flow % Power 50 50 90 10 Fuel per engine, 1320 1320 2180 335 lbmSummation of 2640 lbm 2515 lbm Engine Fuel Used

In various embodiments, and with reference to FIGS. 5A and 5B, anintermediate rated power mode of the multi-engine system 100 isprovided. In the intermediate rated power mode, both gas turbine engines110, 120 may be generally configured to operate on the modifiedoperating line 355 and may both operate at an increased power rating.That is, as schematically depicted and represented by arrows 519 and 529in FIG. 5A, and as it corresponds to the x-axis of the graph of FIG. 5B,both the first gas turbine engine 110 and the second gas turbine engine120 may be configured to operate with a total compressor flow that issubstantially at 100% of a compressor inlet corrected flow capacity ofthe respective compressors. In such a configuration, both the firstcrossover cooling network 131 and the second crossover cooling network132 are transferring inter-engine cooling flow, and thus the firstcrossover airflow flowing through the first crossover cooling network131 may be between about 5% and about 20%, as mentioned above, of thefirst total compressor flow through the first compressor 111 while thesecond crossover cooling network 132 may similarly be between about 5%and about 20% of the second total compressor flow through the secondcompressor 121. Thus, the multi-engine system 100 may operate withbi-directional crossover flow, which may improve the operatingefficiencies of the multi-engine system 100.

In various embodiments, and with reference to FIGS. 6A and 6B, aone-engine-inoperable mode of the multi-engine system 100 is provided.In the one-engine-inoperable mode, for example, the first gas turbineengine 110 is operating at a contingency rated power while the secondgas turbine engine 120 is inoperable. That is, as schematically depictedand represented by arrow 619 in FIG. 6A, and as it corresponds to thex-axis of the graph of FIG. 6B, the first gas turbine engine 110 may beconfigured to operate with a total compressor flow that is substantiallyabove 100% (e.g., 105%) of the compressor inlet corrected flow capacityof the first compressor 111. Said differently, the first gas turbineengine 110, as mentioned above, may have been designed to have extracooling flow and may thus be well-suited to operate at comparativelyhigher contingency power ratings because the excess cooling flow thatwould be routed inter-engine via the crossover networks during normaloperating conditions may be used by the first gas turbine engine 110 atthis overpowered, contingency power during a one-engine-inoperable mode,according to various embodiments.

In various embodiments, crossover flow may cease during operating in theone-engine-inoperable mode. However, in various embodiments, the firstcrossover airflow may continue to flow through the first crossovercooling network 131, thereby improving the possibility for futurerestoration of operation of the second gas turbine engine 120.Accordingly, in various embodiments, the first crossover airflow is 10%of a first total compressor flow through the first compressor 111 whilethe second crossover airflow may be 0.

Other operating modes may be utilized with the crossover coolingconfiguration disclosed herein. Accordingly, the exemplary operatingmodes described herein is not intended as an exhaustive list of theoperating modes of the multi-engine system 100.

In various embodiments, and with reference to FIG. 7, method 790 ofoperating a multi-engine system, such as system 100, is provided. Themethod 790 may include, in response to the multi-engine system operatingin an asymmetric cruise mode, routing a portion of a total compressorflow from a first gas turbine engine through a crossover cooling networkto a second gas turbine engine at step 792. The method 790 may furtherinclude operating the second gas turbine engine at a fuel rate that islower than a self-sustaining idle fuel rate of the second gas turbineengine at step 794. Accordingly, the operation of the second gas turbineengine at step 794 comprises operating said engine at a fuel consumptionrate that is lower than would otherwise be possible if it were not forthe supplemented, crossover airflow from the first gas turbine engine.In various embodiments, during operation of the multi-engine system inthe asymmetric cruise mode, the overall power generation may be the sameas, but the overall fuel consumption may be lower than, if the two gasturbine engines were operating under a symmetric cruise mode (e.g., aconventional cruise mode).

Benefits, other advantages, and solutions to problems have beendescribed herein with regard to specific embodiments. Furthermore, theconnecting lines shown in the various figures contained herein areintended to represent exemplary functional relationships and/or physicalcouplings between the various elements. It should be noted that manyalternative or additional functional relationships or physicalconnections may be present in a practical system. However, the benefits,advantages, solutions to problems, and any elements that may cause anybenefit, advantage, or solution to occur or become more pronounced arenot to be construed as critical, required, or essential features orelements of the disclosure.

The scope of the disclosure is accordingly to be limited by nothingother than the appended claims, in which reference to an element in thesingular is not intended to mean “one and only one” unless explicitly sostated, but rather “one or more.” It is to be understood that unlessspecifically stated otherwise, references to “a,” “an,” and/or “the” mayinclude one or more than one and that reference to an item in thesingular may also include the item in the plural. All ranges and ratiolimits disclosed herein may be combined.

Moreover, where a phrase similar to “at least one of A, B, and C” isused in the claims, it is intended that the phrase be interpreted tomean that A alone may be present in an embodiment, B alone may bepresent in an embodiment, C alone may be present in an embodiment, orthat any combination of the elements A, B and C may be present in asingle embodiment; for example, A and B, A and C, B and C, or A and Band C. Different cross-hatching is used throughout the figures to denotedifferent parts but not necessarily to denote the same or differentmaterials.

The steps recited in any of the method or process descriptions may beexecuted in any order and are not necessarily limited to the orderpresented. Furthermore, any reference to singular includes pluralembodiments, and any reference to more than one component or step mayinclude a singular embodiment or step. Elements and steps in the figuresare illustrated for simplicity and clarity and have not necessarily beenrendered according to any particular sequence. For example, steps thatmay be performed concurrently or in different order are illustrated inthe figures to help to improve understanding of embodiments of thepresent disclosure.

Any reference to attached, fixed, connected or the like may includepermanent, removable, temporary, partial, full and/or any other possibleattachment option. Additionally, any reference to without contact (orsimilar phrases) may also include reduced contact or minimal contact.Surface shading lines may be used throughout the figures to denotedifferent parts or areas but not necessarily to denote the same ordifferent materials. In some cases, reference coordinates may bespecific to each figure.

Systems, methods and apparatus are provided herein. In the detaileddescription herein, references to “one embodiment,” “an embodiment,”“various embodiments,” etc., indicate that the embodiment described mayinclude a particular feature, structure, or characteristic, but everyembodiment may not necessarily include the particular feature,structure, or characteristic. Moreover, such phrases are not necessarilyreferring to the same embodiment. Further, when a particular feature,structure, or characteristic is described in connection with anembodiment, it is submitted that it is within the knowledge of oneskilled in the art to affect such feature, structure, or characteristicin connection with other embodiments whether or not explicitlydescribed. After reading the description, it will be apparent to oneskilled in the relevant art(s) how to implement the disclosure inalternative embodiments.

Furthermore, no element, component, or method step in the presentdisclosure is intended to be dedicated to the public regardless ofwhether the element, component, or method step is explicitly recited inthe claims. No claim element is intended to invoke 35 U.S.C. 112(f)unless the element is expressly recited using the phrase “means for.” Asused herein, the terms “comprises,” “comprising,” or any other variationthereof, are intended to cover a non-exclusive inclusion, such that aprocess, method, article, or apparatus that comprises a list of elementsdoes not include only those elements but may include other elements notexpressly listed or inherent to such process, method, article, orapparatus.

What is claimed is:
 1. A multi-engine system comprising: a first gasturbine engine comprising a first compressor, a first turbine, and afirst power turbine, the first compressor and the first turbine of thefirst gas turbine engine each configured to rotate about a first enginecentral longitudinal axis of the first gas turbine engine; a second gasturbine engine comprising a second compressor, a second turbine, and asecond power turbine, the second compressor and the second turbine ofthe second gas turbine engine each configured to rotate about a secondengine central longitudinal axis of the second gas turbine engine; amain rotor gearbox mechanically coupled to both the first power turbineand the second power turbine; and a first crossover cooling network forrouting a first crossover airflow from the first compressor of the firstgas turbine engine to the second turbine of the second gas turbineengine, wherein the first gas turbine engine and the second gas turbineengine are identical engines.
 2. A multi-engine system comprising: afirst gas turbine engine comprising a first compressor and a firstturbine, the first compressor and the first turbine of the first gasturbine engine each configured to rotate about a first engine centrallongitudinal axis of the first gas turbine engine; a second gas turbineengine comprising a second compressor and a second turbine, the secondcompressor and the second turbine of the second gas turbine engine eachconfigured to rotate about a second engine central longitudinal axis ofthe second gas turbine engine; a first crossover cooling networkconfigured to route a first crossover airflow from the first compressorof the first gas turbine engine to the second turbine of the second gasturbine engine; and a second crossover cooling network configured toroute a second crossover airflow from the second compressor of thesecond gas turbine engine to the first turbine of the first gas turbineengine.
 3. The multi-engine system of claim 2, wherein the multi-enginesystem is configured to operate in a bi-directional crossover mode withboth the first crossover cooling network and the second crossovercooling network having respective crossover flows flowing therethrough.4. The multi-engine system of claim 2, wherein: the first crossovercooling network extends from the first compressor to a second vane rowof the second turbine; and the second crossover cooling network extendsfrom the second compressor to a first vane row of the first turbine. 5.The multi-engine system of claim 4, wherein: the second vane row is aforward-most vane row of the second turbine; and the first vane row is aforward-most vane row of the first turbine.
 6. The multi-engine systemof claim 5, wherein: the forward-most vane row of the second turbinecomprises a plurality of second vanes, wherein each second vane of theplurality of second vanes defines a second leading edge chamber and asecond body chamber aft of the second leading edge chamber, wherein thefirst crossover cooling network is configured to route the firstcrossover airflow to the second body chamber; and the forward-most vanerow of the first turbine comprises a plurality of first vanes, whereineach first vane of the plurality of first vanes defines a first leadingedge chamber and a first body chamber aft of the first leading edgechamber, wherein the second crossover cooling network is configured toroute the second crossover airflow to the first body chamber.
 7. Themulti-engine system of claim 6, wherein: the first gas turbine enginecomprises a first intra-engine cooling network configured to route afirst resident airflow from forward of the first turbine to the firstleading edge chamber of the plurality of first vanes of the forward-mostvane row of the first turbine; and the second gas turbine enginecomprises a second intra-engine cooling network configured to route asecond resident airflow from forward of the second turbine to the secondleading edge chamber of the plurality of second vanes of theforward-most vane row of the second turbine.
 8. The multi-engine systemof claim 2, wherein in response to the multi-engine system operating inan intermediate rated power mode: the first crossover airflow is between5% and 20% of a first total compressor flow through the firstcompressor; and the second crossover airflow is between 5% and 20% of asecond total compressor flow through the second compressor.
 9. Themulti-engine system of claim 8, wherein: the first crossover airflow is10% of the first total compressor flow through the first compressor; andthe second crossover airflow is 10% of the second total compressor flowthrough the second compressor.
 10. The multi-engine system of claim 8,wherein: the first total compressor flow is 100% of a first compressorinlet corrected flow capacity of the first compressor; the second totalcompressor flow is 100% of a second compressor inlet corrected flowcapacity of the second compressor.
 11. The multi-engine system of claim2, wherein in response to the multi-engine system operating in anasymmetric cruise mode: the first crossover airflow is between 5% and20% of a first total compressor flow through the first compressor; andthe second crossover airflow is 0% of a second total compressor flowthrough the second compressor.
 12. The multi-engine system of claim 11,wherein: the first crossover airflow is 10% of the first totalcompressor flow through the first compressor; and the second crossoverairflow is 0% of the second total compressor flow through the secondcompressor.
 13. The multi-engine system of claim 11, wherein: the firsttotal compressor flow is between 90% and 100% of a first compressorinlet corrected flow capacity of the first compressor; and the secondtotal compressor flow is less than or equal to 40% of a secondcompressor inlet corrected flow capacity of the second compressor. 14.The multi-engine system of claim 2, wherein the first gas turbine engineand the second gas turbine engine are identical engines.
 15. Themulti-engine system of claim 14, wherein an electric load, via agenerator, is applied to the second gas turbine engine.
 16. Themulti-engine system of claim 2, wherein in response to the multi-enginesystem operating in a one-engine-inoperable mode: the first crossoverairflow is 10% of a first total compressor flow through the firstcompressor; and the second crossover airflow is 0% of a second totalcompressor flow through the second compressor.
 17. The multi-enginesystem of claim 16, wherein: the first total compressor flow is 105% ofa first compressor inlet corrected flow capacity of the firstcompressor; and the second total compressor flow is 0% of a secondcompressor inlet corrected flow capacity of the second compressor. 18.The multi-engine system of claim 2, wherein: the first crossover coolingnetwork comprises at least one of a first check valve configured toprevent backflow of the first crossover airflow and a first controlledvalve configured to control the first crossover airflow; and the secondcrossover cooling network comprises at least one of a second check valveconfigured to prevent backflow of the second crossover airflow and asecond controlled valve configured to control the second crossoverairflow.
 19. A multi-engine rotorcraft comprising: a first gas turbineengine comprising a first compressor, a first turbine, and a first powerturbine; a second gas turbine engine comprising a second compressor, asecond turbine, and a second power turbine; a main rotor gearboxmechanically coupled to both the first power turbine and the secondpower turbine; a first crossover cooling network configured to route afirst crossover airflow from the first compressor of the first gasturbine engine to a second vane row of the second turbine of the secondgas turbine engine; and a second crossover cooling network configured toroute a second crossover airflow from the second compressor of thesecond gas turbine engine to a first vane row of the first turbine ofthe first gas turbine engine.